Referring to FIG. 1, an aircraft gas turbine engine 1 can have a multistage, rotating air compressor 10 that includes multiple stages of rotor blades 12 and stator vanes 19, a combustion system 3 to burn the fuel, and a multistage turbine 2 to supply the power to the compressor 10 and to direct the gas flow to the engine exit nozzle which provides the thrust force to propel the airplane (not shown). Gas turbine engine fuel efficiency is determined by the compressor pressure ratio and the turbine inlet temperature. Higher pressure ratios generally result in higher compressor exit temperatures. The compressor pressure ratio is limited by the maximum temperature that the materials of the combustion system components (including rotor blades and stator vanes) can withstand, and the maximum turbine inlet temperature is limited by the turbine nozzle vane materials and the effectiveness of the turbine cooling system.
Several stationary gas turbine engines have been designed with low pressure and high pressure compressors and with very large and heavy inter-cooling systems that cool all of the airflow between and within the compressors to achieve a lower compressor exit temperature. These engines have higher pressure ratios, which increases the fuel efficiency of the engine, without higher compressor exit temperatures.
The advantages of inter-cooling in the compressor have been recognized for many years, but no practical light weight designs for inter-cooling compressors for aircraft engines have been proposed or developed.